Previous composite insulating structures intended for use in such applications as on orbital reentry vehicles, such as the Space Shuttle Orbiter, have consisted of a coating in combination with a low density insulation substrate. A borosilicate glass, Reaction Cured Glass (RCG), was chosen as the coating for the silica type of Reusable Surface Insulation (RSI) previously selected as the heat shield for the Orbiter, as disclosed in U.S. Pat. No. 4,093,771 issued Jun. 6, 1978 to Goldstein et al. This coating was prepared by blending an emittance agent, silicon tetraboride, with a specially prepared borosilicate glass powder, composed of approximately 94% by weight silica and 6% by weight boron oxide, and an ethanol carrier in a ball mill. The resulting slurry was optimized to limit penetration of the slurry into the low density insulation and limit sagging during the subsequent sintering operation. After drying, the coating was sintered (fused) to a “theoretical” density of 2.2 g/cc at 1220° C. (2225° F.). The “as fired” RCG coating weighs 0.07 g/sq cm (0.15 lbs/sq ft) and is approximately 0.3 mm (0.013 in) thick and has been applied to advanced insulation systems.
Current passive systems being proposed for use as leading edge thermal protection systems on future vehicles include hot structure, heat sink, and transpiration cooled technologies. Hot structures such as reinforced carbon-carbon are being used on the Orbiter, e.g., on the wing leading edge and nose cap, and silicon carbide systems have been proposed for other reentry vehicles such as the X-33 and X-38, etc.
The systems used on the Orbiter for the wing leading edge and nose cap typically operate at temperatures below 2700° F. (1480° C.) during Earth entry. This system, if breached, has resulted in the loss of the vehicle during Earth entry. Also, these systems are heavier, orders-of-magnitude more expensive, and require much longer lead times for delivery than a system made using rigid fibrous insulation. Other lighter weight, less expensive alternatives including coated fibrous insulation systems (such as RCG on LI-900, a low-density fibrous silica glass structure) are susceptible to excessive surface recession and/or impact damage during launch and/or landing.
The impact resistance of the fibrous insulation systems was substantially improved by the development of the Toughened Uni-Piece Fibrous Insulation (TUFI) as disclosed in U.S. Pat. No. 5,079,082 issued Jan. 7, 1992 to Leiser et al. This material represented the first family of lightweight inexpensive graded thermal protection materials. The emittance agent was molybdenum disilicide. It included a borosilicate glass matrix and a processing aid (silicon hexaboride) that permitted sintering at 2225° F. (1220° C.). TUFI is used in selected areas on the Orbiter where the temperatures of operation are much lower than its capability of 2600° F. (1425° C.). The use of TUFI-like systems as a leading edge requires still higher temperature capability.
The specific disadvantages of the prior art depend upon the type of leading edge chosen. For a structural type leading edge, the major disadvantages are weight; the complexity of designing a hot structure that must perform under load and expand from room temperature to high temperature while maintaining structural integrity and while remaining attached to the vehicle; and the cost and time required to produce appropriate parts. For an insulating leading edge the disadvantage has been the relatively limited temperature capability of materials that has made it, under most circumstances, an impracticable choice due to the unacceptable limitations it imposes on the resultant atmospheric entry vehicles.
For the reasons stated above, and for other reasons stated below that will become apparent to those skilled in the art upon reading and understanding the present specification, there is a need in the art for alternative insulating structures.